Torque generating turbines are rotary engines that extract energy from a fluid flow, e.g. steam flow and/or combustion gas flow, and convert it into useful work. Typically, torque-generating turbines are used to drive a generator to produce electric energy. Usually, a torque-generating turbine comprises a stator and a rotor having one or several turbine stages.
In a turbine stage fluid, e.g. steam or exhaust gas, is directed onto the rotor by the fixed vanes of the stator. It leaves the stator as a jet that fills the entire circumference of the rotor. The flow then changes direction and increases its speed relative to the speed of the blades. A pressure drop occurs across both the vane and the blade, with flow accelerating through the vane and decelerating through the blade, with no significant in flow velocity before and after the stage (vane inlet and blade outlet) but with a decrease in both pressure and temperature, reflecting the work performed in the driving of the rotor.
Torque-generating gas turbines additionally comprise a compressor and a combustion chamber. The compressor having a stator and a rotor like the turbine typically has several stages as described in the paragraph hereinbefore. However, in the compressor both pressure and temperature increases from the first to the last stage, reflecting the work performed by the turbine in the driving of the rotor of the compressor. The compressor compresses ambient air before it is mixed with fuel in the combustion chamber to generate a burning mixture. The exhaust gas is then fed into the turbine in the narrower sense where it expands. After a portion of the turbine, e.g. after 2 stages, an intermediate diffuser may be provided for reducing the speed of the fluid flow such that its flow adapts smoothly to the geometry of the downstream turbine stages. After the last turbine stage there is a diffuser that decelerates the flow as much as possible to minimize the pressure drop and thereby improve the performance of the turbine and the gas turbine as a whole before it is exhausted through the stack. The compressor also has a diffuser after its last stage to minimize the pressure drop. The rotating rotor of the turbine then drives the compressor and for example a generator.
The junction of the vanes of the stators of either the compressor stages or the turbine stages with the outer or inner surfaces may be sources of horseshoe vortexes and may thus impede, that the fluid flow is optimally redirected onto the rotor.
EP 1 674 664 A2 relates to an exit guide vane array for thrust generating military aircraft engines, which includes a set of guide vanes having a solidity and defining fluid flow passages with a chordwisely converging forward portion. The high solidity and convergent passage portion resist fluid separation. The vanes cooperate with each other to restrict an observer's line of sight of planes upstream of the van array.
In EP 0798447 B1 a gas turbine guide vane with an enlarged filled radius near the end wall has been proposed to reduce the amount of drag generating vortexes. Furthermore, WO 00/61918 suggests enlarging the leading edge radius near the end walls. In WO 2010/063271 it has additionally been proposed to enlarge the cross sectional area of struts supporting annular channels connecting subsequent turbine stages of a thrust generating aircraft gas turbine.
The fluid flow velocity after the rotor may still be significant. Accordingly, turbines are provided after the last rotor with an exhaust diffuser to slow down the fluid flow and thereby enhance pressure recovery.
The exhaust diffuser typically comprises an inner member having an outer surface and an outer member having an inner surface, wherein the inner member and the outer member form an annular channel. The inner member has to be supported within the outer member. Accordingly, the exhaust diffuser comprises struts extending essentially radially from the outer surface of the inner member to the inner surface of the outer member. Typically, the struts have a prismatic design with a cross section that does not change across the span from the outer surface of the inner member to the inner surface of the outer member.
The cross section of the strut typically has the form of an airfoil, i.e. it is shaped aerodynamically to reduce the drag induced by the strut as far as possible. The strut often has more than one function. Maintaining the distance between the inner member and outer member, transfer forces from the rotor via the bearing to the casing, provide inspection access to the area inside the inner member, provide passage for instrumentation wiring, supply and drain routes of air and lubrication oil to bearing.
Several parameters may characterize an airfoil: —chord line, —chord, —thickness, —mean camber line, —camber, —leading edge radius
The chord line is shortest line connecting leading edge and trailing edge of an airfoil. Accordingly, the chord denotes its length. Unless otherwise stated, further dimensions of an airfoil are always given relative to the chord. The thickness of an airfoil indicates the maximum extension perpendicular to the chord line. The mean camber line is the line connecting the points midway between the upper surface and the lower surface of an airfoil along the chord line. The camber expresses the maximum distance between the camber line and the chord line. For symmetric airfoils chord line and camber line are identical and thus the camber is zero, of course. The leading edge radius is the radius, which may be fitted to the leading edge of the airfoil.
A fluid flow component at the inlet of the exhaust diffuser may comprise a component transversal to the exhaust diffuser, i.e. may enclose a flow angle different from zero with an axis of the annular channel of the exhaust diffuser. The flow angle may depend on the capacity of the torque-generating turbine or its operating point, e.g. load or speed, and the radial distance of the fluid flow component from the axis of the annular channel of the exhaust diffuser.
The flow angle significantly affects pressure recovery of an exhaust diffuser. It has been shown that an exhaust diffuser may show good performance up to a flow angle of 15 degree whereas steep losses occur thereafter.
The angle of incidence of an airfoil of conventional prismatic struts, i.e. the angle between the chord line and the axis of the exhaust diffuser, is typically selected to be zero or to be equal to the mean inlet flow angle. Hence, the angle of attack, i.e. the angle between the fluid flow and the chord line of the airfoil may therefore vary across the span of the prismatic strut. If the angle of attack becomes too high or too low flow separation may occur and large regions of low momentum fluid may be generated. These may lead to blockages and endanger pressure recovery of the exhaust diffuser. Accordingly, airfoils supporting a large range of angle of attack, in particular airfoils having a large leading edge radius have been used hereinbefore. However, airfoils supporting a large range of angle of attack may have a higher drag coefficient when the angle of attack is low thus rendering the exhaust diffuser less efficient.
US 2011/0052373 A1 proposes a supporting strut comprising a channel from the pressure side to the suction side of an airfoil to avoid flow separation. However, this supporting strut has a significant cross-section when the angle of attack is low going along with a high drag.
US 2009/0324400 A1 suggests a gas turbine for use in subsonic flight wherein a supporting strut for the thrust generating nozzle of the gas turbine is provided with a channel to reduce strut wake loss.
There may be a need for a more efficient exhaust diffuser being at the same time less prone to blockages.